Modulated turbine component cooling

ABSTRACT

Features and methods for modulating a flow of cooling fluid to gas turbine engine components are provided. In one embodiment, an airfoil is provided having a flow modulation insert for modulating a flow of cooling fluid received in a cavity of a body of the airfoil. In another embodiment, a shroud is provided comprising a cooling channel for a flow of cooling fluid and an insert that varies in position to modulate the flow of cooling fluid through the cooling channel. In yet another embodiment, a method for operating a gas turbine engine having a cooling circuit for cooling one or more components of the gas turbine engine comprises increasing power provided to the engine and decreasing power provided to the engine to modulate a position of a flow modulation insert located in the cooling circuit and thereby modulate the flow of cooling fluid through the cooling circuit.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of and claims priority to U.S.application Ser. No. 15/231,846, filed Aug. 9, 2016, the contents ofwhich are incorporated herein in their entirety by reference thereto.

FIELD OF THE INVENTION

The present subject matter relates generally to gas turbine engines andparticularly to features for cooling components of gas turbine engines.More particularly, the present subject matter relates to modulating aflow of cooling fluid through components of gas turbine engines.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gasesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

The combustion gases have temperatures that can be detrimental to enginecomponents, e.g., by causing material failures or the like. Typically, aflow of cooling fluid is provided to engine components exposed to thecombustion gases to reduce or mitigate negative impacts of thecombustion gas temperatures. However, the cooling flow may not berequired or desirable during all operating conditions. For example,specific fuel consumption (SFC) may be reduced by reducing orrestricting the flow of cooling fluid during non-high power cycleconditions, such as a cruise operating condition. Moreover, typicalsystems for providing cooling flow rely on a network of valves and pipesthat can increase the engine weight, thereby increasing fuelconsumption, without providing a benefit that sufficiently offsets thenegative impacts of increased engine weight.

Therefore, improved cooling features that overcome one or moredisadvantages of existing components and systems would be desirable. Inparticular, an insert for a turbine section component that modulates aflow of cooling fluid based on changes in temperature would bebeneficial. More particularly, such an insert that shifts position whenexposed to increased temperatures to allow an increase in cooling flowwould be desirable. Further, such an insert that utilizes a differencein a coefficient of thermal expansion of the insert and the turbinesection component to modulate the cooling flow would be advantageous.Additionally or alternatively, such an insert that utilizes a shapememory alloy to modulate the cooling flow would be useful.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, an airfoil for agas turbine engine is provided. The airfoil comprises a body having apressure side and a suction side. The body defines a cavity for receiptof a flow of cooling fluid. The airfoil further includes a flowmodulation insert positioned within the cavity. The flow modulationinsert defines a plurality of apertures permitting the cooling fluid toflow therethrough. The airfoil also includes an attachment element thatvaries in position between a first position and a second position. Theflow modulation insert is coupled to the attachment element such that alocation of the plurality of apertures of the flow modulation insert isvariable to modulate the flow of cooling fluid through the apertures ofthe flow modulation insert.

In another exemplary embodiment of the present disclosure, a shroud fora gas turbine engine is provided. The shroud comprises a cooling channelfor a flow of cooling fluid. The cooling channel has an inlet defined ina forward end of the shroud and an outlet defined within the shroudadjacent an internal cooling passage of the shroud. The shroud alsoincludes an insert that varies in position between a first position anda second position based on changes in an environmental temperature ofthe insert. The first position blocks the cooling channel to reduce theflow of cooling fluid therethrough, and the second position unblocks thecooling channel to increase the flow of cooling fluid therethrough.

In a further exemplary embodiment of the present disclosure, a methodfor operating a gas turbine engine is provided. The gas turbine engineincludes a cooling circuit for providing a flow of cooling fluid to oneor more components of the gas turbine engine. The method comprisesincreasing power provided to the gas turbine engine and decreasing powerprovided to the gas turbine engine, where increasing and decreasing thepower provided to the gas turbine engine modulates a position of a flowmodulation insert located in the cooling circuit to modulate the flow ofcooling fluid through the cooling circuit.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-sectional view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a close-up, side view of a portion of a turbine sectionof the exemplary gas turbine engine of FIG. 1 .

FIG. 3A provides a cross-section view of a shroud of the turbine sectionof FIG. 2 having an insert in a first position, according to anexemplary embodiment of the present subject matter.

FIG. 3B provides the cross-section view of the shroud of FIG. 3A havingthe insert in a second position, according to an exemplary embodiment ofthe present subject matter.

FIG. 3C provides a cross-section view of a shroud of the turbine sectionof FIG. 2 having an insert in a first position, according to anexemplary embodiment of the present subject matter.

FIG. 3D provides the cross-section view of the shroud of FIG. 3C havingthe insert in a second position, according to an exemplary embodiment ofthe present subject matter.

FIG. 4 provides a forward end view of a portion of the shroud of FIG. 2, according to an exemplary embodiment of the present subject matter.

FIG. 5 provides an axial cross-section view of an airfoil of a turbinenozzle, according to an exemplary embodiment of the present subjectmatter.

FIG. 6A provides a radial cross-section view of the airfoil of FIG. 5with a flow modulation insert in a first position, according to anexemplary embodiment of the present subject matter.

FIG. 6B provides the cross-section view of the airfoil of FIG. 6A withthe flow modulation insert in a second position, according to anexemplary embodiment of the present subject matter.

FIG. 7 provides an axial cross-section view of an airfoil of a turbinenozzle, according to an exemplary embodiment of the present subjectmatter.

FIG. 8A provides a radial cross-section view of the airfoil of FIG. 7with a flow modulation insert in a first position, according to anexemplary embodiment of the present subject matter.

FIG. 8B provides the cross-section view of the airfoil of FIG. 8A withthe flow modulation insert in a second position, according to anexemplary embodiment of the present subject matter.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1 , the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1 , the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. Fan blades 40 and disk 42 are togetherrotatable about the longitudinal axis 12 by LP shaft 36. In someembodiments, fan 38, including fan blades 40 and disk 42, may berotatable across a power gear box that includes a plurality of gears forstepping down the rotational speed of the LP shaft 36 to a moreefficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1 , disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

In some embodiments, components of turbofan engine 10, particularlycomponents within hot gas path 78 such as components within thecombustion section 26 or downstream thereof, may comprise a ceramicmatrix composite (CMC) material, which is a non-metallic material havinghigh temperature capability. In general, turbine performance andefficiency may be improved by increased combustion gas temperatures;therefore, non-traditional high temperature materials, such as CMCmaterials, are more commonly being used for various components withingas turbine engines, including components within the flow path of thecombustion gases. Exemplary CMC materials utilized for gas turbineengine components may include silicon carbide (SiC), silicon, silica, oralumina matrix materials and combinations thereof. Ceramic fibers may beembedded within the matrix, such as oxidation stable reinforcing fibersincluding monofilaments like sapphire and silicon carbide (e.g.,Textron's SCS-6), as well as rovings and yarn including silicon carbide(e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and DowCorning's SYLRAIVIIC®), alumina silicates (e.g., Nextel's 440 and 480),and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), andoptionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, andcombinations thereof) and inorganic fillers (e.g., pyrophyllite,wollastonite, mica, talc, kyanite, and montmorillonite). For example, incertain embodiments, bundles of the fibers, which may include a ceramicrefractory material coating, are formed as a reinforced tape, such as aunidirectional reinforced tape. A plurality of the tapes may be laid uptogether (e.g., as plies) to form a preform component. The bundles offibers may be impregnated with a slurry composition prior to forming thepreform or after formation of the preform. The preform may then undergothermal processing, such as a cure or burn-out to yield a high charresidue in the preform, and subsequent chemical processing, such asmelt-infiltration with silicon, to arrive at a component formed of a CMCmaterial having a desired chemical composition. In other embodiments,the CMC material may be formed as, e.g., a carbon fiber cloth ratherthan as a tape.

As stated, components comprising a CMC material may be used within thehot gas path 78, such as within the combustion and/or turbine sectionsof engine 10. However, CMC components may be used in other sections aswell, such as the compressor and/or fan sections. As a particularexample described in greater detail below, a turbine nozzle or a turbineshroud may be formed from a CMC material to better withstand the heat ofthe combustion gases, as well as to withstand increased combustion gastemperatures.

It will be appreciated that, although described with respect to turbofanengine 10 having core turbine engine 16, the present subject matter maybe applicable to other types of turbomachinery. For example, the presentsubject matter may be suitable for use with or in turboprops,turboshafts, turbojets, industrial and marine gas turbine engines,and/or auxiliary power units.

Referring now to FIG. 2 , a schematic view is provided of the HP turbine28 of the turbine section of core turbine engine 16, which is locateddownstream from combustion section 26. The combustion section 26generally includes a combustor defining a combustion chamber 80; amixture of fuel and air is combusted within the combustion chamber togenerate a flow of combustion gases 66 therethrough. Downstream of thecombustion section 26, the HP turbine 28 includes a first turbine nozzlestage 100, a turbine blade stage 102, and a second turbine nozzle stage104, each configured to direct a flow of combustion gasses therethrough.Notably, the first turbine nozzle stage 100 is located immediatelydownstream from the combustion section 26, and thus may also be referredto as a combustor discharge nozzle stage. Further, it should beunderstood that, although only two nozzle stages and one blade stage areillustrated in FIG. 2 , HP turbine 28 may include a plurality of nozzleand blade stages. Moreover, LP turbine 30 likewise may be configuredwith a plurality of turbine nozzle and turbine blade stages, and eachstage of HP turbine 28 and LP turbine 30 may comprise a plurality ofturbine components that define and/or are positioned within the hot gaspath 78 through which the combustion gases flow.

The first turbine nozzle stage 100 includes a plurality of turbinenozzle sections 106 spaced along a circumferential direction C (FIG. 4). Each first turbine nozzle section 106 forming the first turbinenozzle stage 100 includes a first stage turbine nozzle 108 positionedwithin the hot gas path 78. Further, each nozzle section 106 includes aninner band segment 110 defining an inner wall of the nozzle section 106and an outer band segment 112 defining an outer wall of the nozzlesection 106, with nozzle 108 extending generally along the radialdirection R from inner band segment 110 to outer band segment 112.Together, the plurality of first turbine nozzle sections 106 define thefirst turbine nozzle stage 100, with an inner band defined by theplurality of inner band segments 110 of nozzle sections 106, an outerband defined by the plurality of outer band segments 112 of nozzlesections 106, and a plurality of nozzles 108 extending from the innerband to the outer band. In some embodiments, the inner band and/or theouter band may be formed as a single continuous component rather thanfrom a plurality of inner and outer band segments 110, 112. As such, inappropriate embodiments, first turbine nozzle stage 100 may be formedfrom an inner band and an outer band with a plurality of nozzles 108extending therebetween, rather than from a plurality of nozzle segmentscomprising an inner band segment 110, an outer band segment 112, and oneor more nozzles 108.

Located immediately downstream of the first turbine nozzle stage 100 andimmediately upstream of the second turbine nozzle stage 104, the HPturbine 28 includes a first stage 102 of turbine rotor blades 116. Firststage 102 of turbine rotor blades 116 includes a plurality of turbinerotor blades 116 spaced along the circumferential direction C and afirst stage rotor 118. Each turbine rotor blade 116 is attached to thefirst stage rotor 118. Although not depicted, the first stage turbinerotor 118 is, in turn, connected to the HP shaft 34 (FIG. 1 ). In suchmanner, the turbine rotor blades 116 may extract kinetic energy from theflow of combustion gases through the hot gas path 78 defined by the HPturbine 28 as rotational energy applied to the HP shaft 34. Core gasturbine engine 16 additionally includes a shroud 120 exposed to and atleast partially defining hot gas path 78. Shroud 120 is described ingreater detail below.

Similar to the plurality of nozzle sections 106 forming the firstturbine nozzle stages 100, a radially inner portion of each turbinerotor blade 116 includes a wall or platform 122. Additionally, eachturbine rotor blade 116 includes a tip 124 at a radially outer portionof the blade. Shroud 120 may be positioned radially adjacent yet spacedapart from the blade tips 124 such that shroud 120 defines an outer wallof the rotor blade stage 102. Further, shroud 120 includes a radiallyouter surface 126 opposite a radially inner surface 128 that is exposedto and at least in part defines the hot gas path 78. As shown in FIG. 2, shroud 120 is tightly configured relative to the blades 116 so thatthe shroud 120 defines an outer radial flow path boundary for the hotcombustion gas flowing through the turbine 16. Shroud 120 generallyforms a ring or shroud around the first stage 102 of turbine rotorblades 116, i.e., shroud 120 extends circumferentially about thelongitudinal engine axis 12 proximate the turbine rotor blade stage 102.As such, in the depicted exemplary embodiment, shroud 120 is an annularshroud that extends circumferentially around first stage 102 of turbinerotor blades 124. In some embodiments, shroud 120 may be formed as acontinuous, unitary, or seamless ring. However, in other embodimentsshroud 120 may be formed from a plurality of shroud segments thattogether form shroud 120. Thus, in such embodiments, shroud 120 mayinclude a plurality of shroud segments positioned next to one anotheralong the circumferential direction C to form generally annular shroud120 around first turbine rotor blade stage 102.

As further depicted in FIG. 2 , the second turbine nozzle stage 104includes a plurality of second turbine nozzle sections 130 spaced alonga circumferential direction C (FIG. 4 ). Each second turbine nozzlesection 130 forming the second turbine nozzle stage 104 includes aplurality of second stage turbine nozzles 132 positioned within the hotgas path 78. Moreover, second turbine nozzle stage 104 includes aplurality of inner band segments 134 forming an inner wall of secondturbine nozzle stage 104, as well as a plurality of outer band segments136 forming an outer wall of second turbine nozzle stage 104. Eachsecond stage turbine nozzle 132 extends generally along the radialdirection R from an inner band segment 134 to an outer band segment 136.Similar to the first turbine nozzle stage 100, the plurality of secondturbine nozzle sections 130 together define the second turbine nozzlestage 104, with an inner band defined by the plurality of inner bandsegments 134 of nozzle sections 130, an outer band defined by theplurality of outer band segments 136 of nozzle sections 130, and aplurality of nozzles 132 extending from the inner band to the outerband. In some embodiments, the inner band and/or the outer band may beformed as a single continuous component rather than from a plurality ofinner and outer band segments 134, 136. As such, in appropriateembodiments, second turbine nozzle stage 104 may be formed from an innerband and an outer band with a plurality of nozzles 132 extendingtherebetween, rather than from a plurality of nozzle segments comprisingan inner band segment 134, an outer band segment 136, and one or morenozzles 132.

Further, although described herein with respect to HP turbine 28, thepresent subject matter is not limited to HP turbines but mayadditionally or alternatively be utilized in a similar manner in the lowpressure compressor 22, high pressure compressor 24, LP turbine 30,and/or any other suitable component of turbofan engine 10.

Referring to FIGS. 3A and 3B, cross-sectional views are provided of aportion of shroud 120 according to an exemplary embodiment of thepresent subject matter. As shown in FIG. 3A and FIG. 3B, shroud 120defines a cooling channel 200 for directing a flow of cooling fluidtherethrough. The flow of cooling fluid may be, e.g., a flow ofpressurized air diverted from HP compressor 24. The cooling channel 200has an inlet 202 defined in a forward end of shroud 120 and an outlet204 defined within shroud 120 adjacent an internal cooling passage 206of shroud 120. Thus, the cooling channel 200 provides fluidcommunication from the forward end of the shroud 120 to cooling passage206. More specifically, the inlet 202 of cooling channel 200 has aninlet flow area over which cooling fluid flow is permitted to coolingchannel 200. Similarly, the outlet 204 of cooling channel 200 has anoutlet flow area over which cooling fluid flow is permitted to coolingpassage 206.

As further shown in FIG. 3A, FIG. 3B, and FIG. 4 , an insert 208 ispositioned at the inlet 202 of cooling channel 200. Referringparticularly to FIG. 4 , insert 208 may substantially surround or bepositioned over inlet 202 on a forward end 121 of shroud 120, but insert208 may have other configurations as well. Insert 208 changes positionsin response to temperature changes to modulate the flow of cooling fluidto cooling channel 200. More particularly, insert 208 has a firstposition 210, shown in FIG. 3A, that the insert maintains until athreshold temperature is reached. In the first position 210, insert 208restricts or blocks the flow of cooling fluid to cooling channel 200,i.e., insert 208 in the first position reduces the inlet flow area ofinlet 202. In some embodiments, the first position 210 of insert 208 mayprovide a reduction in inlet flow area of at least about 50%, generallywithin a range of about 50% to about 90%. In an exemplary embodiment,the first position 210 reduces the inlet flow area by about 80% to about85%. In other embodiments, the first position 210 may block essentiallyall cooling fluid flow through inlet 202, i.e., the first position mayprovide an approximately 100% reduction in inlet flow area.

At the threshold temperature, the insert 208 transitions to a secondposition 212, shown in FIG. 3B. In the second position 212, insert 208provides little to no restriction of cooling fluid flow through inlet202 of cooling passage 200. That is, if the insert 208 transitions fromthe first position 210 to the second position 212, the insert 208 opensup the inlet 202 to permit a greater flow of cooling fluid through inlet202 than was permitted when the insert 208 was in the first position210. In some embodiments, the second position 212 of insert 208 mayblock only about 5% to about 20% of the inlet flow area. In an exemplaryembodiment, the second position 212 reduces the inlet flow area by about10% or less. In other embodiments, the second position 212 mayessentially provide no restriction of cooling fluid flow through inlet202, i.e., the second position may permit flow over approximately 100%of the inlet flow area.

As shown in FIGS. 3C and 3D, in other embodiments insert 208 mayadditionally or alternatively be positioned at outlet 204 of coolingchannel 200 to modulate the flow of cooling fluid by variouslyrestricting or not restricting the outlet flow area. Referring to FIG.3C, in the first position 210, insert 208 restricts the outlet flow areaof outlet 204, thereby restricting the cooling fluid flow to coolingpassage 206. As depicted in FIG. 3D, in the second position 212, insert208 unblocks or does not restrict the outlet flow area of outlet 204such that a greater flow of cooling fluid may be provided to coolingpassage 206. Similar the modulation of the inlet flow area, the firstposition 210 of insert 208 may restrict the outlet flow area by at leastabout 50%, generally within a range of about 50% to about 90%. In oneembodiment, the first position 210 reduces the outlet flow area by about80% to about 85%. In other embodiments, the first position 210 may blockessentially all cooling fluid flow through outlet 204, i.e., the firstposition may provide an approximately 100% reduction in outlet flowarea. Further, with respect to the second position 212, insert 208 mayblock only about 5% to about 20% of the outlet flow area. In anexemplary embodiment, the second position 212 reduces the outlet flowarea by about 10% or less. In other embodiments, the second position 212may essentially provide no restriction of cooling fluid flow throughoutlet 204, i.e., the second position may permit flow over approximately100% of the outlet flow area.

It will be appreciated that in either the embodiment of FIGS. 3A and 3Bor the embodiment of FIGS. 3C and 3D, insert 208 may transition betweenfirst position 210 and second position 212 based on changes intemperature. For example, insert 208 may be in the first position 210until the temperature of its surrounding environment rises above athreshold temperature; then, insert 208 transitions to the secondposition 212 in response to the change in temperature to a temperaturegreater than the threshold temperature. However, if the temperature ofthe environment of insert 208 cools to below the threshold temperature,insert 208 returns to the first position 210. That is, insert 208changes from its first position 210 to its second position 212 or viceversa as the temperature of its environment rises above or falls belowthe threshold temperature.

Insert 208 may be made from a material with a relatively largecoefficient of thermal expansion, i.e., a relatively high alphamaterial, or a shape memory alloy such that the insert 208 can expandand contract in response to temperature changes. More particularly, insome embodiments, insert 208 and shroud 120 may be formed from materialshaving different coefficients of thermal expansion, such that thematerial of each component has different thermal growth characteristics.For example, insert 208 may be formed from a metal material and shroud120 from a CMC material; in such embodiments, insert 208 is a high alphamaterial in comparison to the shroud 120. As such, the insert 208expands or grows to a greater extent than the shroud 120 at a giventemperature, i.e., the insert 208 expands or grows more than the shroud120 as the surrounding temperature begins to increase. The high alphamaterial of insert 208 may be selected such that insert 208 expands tothe second position 212 above a certain temperature.

In other embodiments, insert 208 may be formed from a shape memoryalloy. A shape memory alloy can exist in two distinct temperaturedependent crystal structures or phases. The temperature at which a phasechange occurs between the crystal structures is dependent upon thecomposition of the alloy, and the phase change temperature is known asthe transition temperature. For example, one distinct crystal structure,known as martensite, corresponds to a lower temperature and a seconddistinct crystal structure, known as austenite, corresponds to a highertemperature. A two-way shape memory alloy has the ability to recover apreset shape upon heating above the transition temperature and to returnto a certain alternate shape upon cooling below the transitiontemperature. The two-way shape memory alloy may be programmed or trainedthrough a process of mechanical working and heat treatment so that itresponds to temperature changes and/or the transition temperature in apredictable and repeatable manner. Thus, in some embodiments, insert 208may be formed from a shape memory alloy such that insert 208 assumes thefirst position 210 below a threshold or transition temperature andassumes the second position 212 above the threshold or transitiontemperature. In alternate embodiments, the insert 208 may comprise abi-metallic material that responds to the threshold temperature in asimilar manner to a shape memory alloy. In still other embodiments, theinsert 208 may comprise a high temperature shape memory polymer thatresponds to the threshold temperature in a similar manner to a shapememory alloy. In particular embodiments, the insert 208 is constructedof a two-way shape memory alloy such as nickel titanium (NiTi) alloyhaving a phase change or transition temperature within a heat transientof the cooling fluid flowing, e.g., between the compressor section 24and the cooling passages of the turbine section. In such embodiments, atleast a portion of the insert 208 is subjected to a programming processin which the insert 208 assumes the first position 210 in a martensiteor lower temperature configuration and assumes the second position 212in an austenite or higher temperature configuration.

Turning now to FIG. 5 , an axial cross-section view is provided of anairfoil of a turbine nozzle, such as first stage turbine nozzle 108 orsecond stage turbine nozzle 132, according to an exemplary embodiment ofthe present subject matter. Each turbine nozzle includes an airfoil 300that extends within the hot gas path 78 such that the combustion gases66 flow against and around the airfoil. Each airfoil 300 has a body 302that includes a concave pressure side 304 opposite a convex suction side306. Opposite pressure and suction sides 304, 306 of each airfoil 300extend axially between a leading edge 308 and an opposite trailing edge310. Leading edge 308 defines a forward end of airfoil 300, and trailingedge 310 defines an aft end of airfoil 300. Further, pressure andsuction sides 304, 306 of airfoil 300 define an outer surface 312 of theairfoil body 302. Additionally, an inner surface 314 of body 302 definesa cavity 316, which in the depicted embodiment is closer to leading edge308 than trailing edge 310 and mostly forward of an axial midpoint ofthe airfoil, for receiving a flow of cooling fluid, e.g., a flow ofpressurized air diverted from HP compressor 24. The flow of coolingfluid may be directed to one or more portions of airfoil 300, e.g., tocool the airfoil and thereby mitigate the impacts of the temperatures ofthe combustion gases that flow against and around the airfoil 300.

As further depicted in FIG. 5 , a flow director 318 is positioned in thecavity 316, and a flow modulation insert 320 is positioned adjacent theflow director 318. In the embodiment of FIG. 5 , the flow modulationinsert 320 is positioned inward of the flow director 318 such that aspace 322 is defined between inner surface 314 of body 302 and the flowdirector 318. In other embodiments, the flow modulation insert 320 maybe positioned outward of flow director 318 such that the space 322 isdefined between inner surface 314 and flow modulation insert 320.

Referring to FIGS. 6A and 6B, radial cross-section views are provided ofthe airfoil 300 of FIG. 5 , according to an exemplary embodiment of thepresent subject matter. As shown in FIGS. 6A and 6B, the flow director318 defines a plurality of apertures 324 therethrough, e.g., fordirecting the flow of cooling fluid received in cavity 316 to the innersurface 314 of airfoil body 302. Each aperture 324 has an aperture flowarea over which cooling fluid flow may flow toward inner surface 314 ofbody 302. The flow modulation insert 320 also defines a plurality ofapertures 326 therethrough, which may allow the flow of cooling fluid Fto flow through flow modulation insert 320.

Similar to insert 208 described above, flow modulation insert 320changes positions in response to temperature changes to modulate theflow of cooling fluid to body 302. More specifically, flow modulationinsert 320 has a first position 330, shown in FIG. 6A, that the insertmaintains until its environment reaches a threshold temperature. In thefirst position 330, flow modulation insert 320 restricts or blocks theflow of cooling fluid to apertures 324, i.e., flow modulation insert 320in the first position reduces the aperture flow area of each aperture324. In some embodiments, the first position 330 of flow modulationinsert 320 may provide a reduction in aperture flow area of about 50% toabout 90%. In an exemplary embodiment, the first position 330 reducesthe aperture flow area by about 80% to about 85%. In other embodiments,the first position 330 may block essentially all cooling fluid flowthrough apertures 324, i.e., the first position may provide anapproximately 100% reduction in aperture flow area.

At the threshold temperature, the flow modulation insert 320 transitionsto a second position 332, shown in FIG. 6B. In the second position 332,flow modulation insert 320 provides little to no restriction of coolingfluid flow through apertures 324 of flow director 318. That is, if theflow modulation insert 320 transitions from the first position 330 tothe second position 332, the flow modulation insert 320 opens up theapertures 324 to permit a greater flow of cooling fluid F throughapertures 324 than was permitted when the flow modulation insert 320 wasin the first position 330. In some embodiments, the second position 332of flow modulation insert 320 may block only about 5% to about 20% ofthe aperture flow area. In an exemplary embodiment, the second position332 reduces the aperture flow area by about 10% or less. In otherembodiments, the second position 332 may essentially provide norestriction of cooling fluid flow through apertures 324 of flow director318, i.e., the second position may permit flow over approximately 100%of the flow area of apertures 324. Accordingly, the flow modulationinsert 320 may shift positions such that the location of its apertures326 varies with respect to the apertures 324 of flow director 318 tomodulate the flow of cooling fluid to airfoil body 302.

Similar to insert 208 previously described, in some embodiments flowmodulation insert 320 may be made from a material with a largecoefficient of thermal expansion, i.e., a high alpha material, or ashape memory alloy such that the flow modulation insert 320 can expandand contract in response to temperature changes. In an exampleembodiment in which the flow modulation insert 320 is formed from a highalpha material, the insert 320 may be formed from a metal material andthe flow director 318 from a CMC material such that the insert 320expands or grows to a greater degree than the flow director 318 as thetemperature rises. For example, as the temperature increases, theapertures 326 of flow modulation insert 320 may expand incross-sectional area more than apertures 324 of flow director 318 suchthat the flow modulation insert 320 no longer blocks apertures 324 orrestricts flow through apertures 324 to a lesser extent than the insert320 restricts the flow at lower temperatures. In other embodiments, flowmodulation insert 320 may be formed from a shape memory alloy such thatthe insert 320 shifts position with respect to flow director 318 at anincreased temperature to permit greater fluid flow through apertures 324of flow director 318 and thereby provide greater cooling to airfoil body302.

In still other embodiments, rather than forming flow modulation insert320 from a high alpha material or a shape memory alloy, the airfoil 300may include an attachment element 328 coupled to the flow modulationinsert 320 that changes position in response to changes in temperatureand thereby shifts the position of insert 320 in response to changes intemperature. As shown in FIG. 6A, the attachment element 328 takes on afirst configuration to place the flow modulation insert 320 in the firstposition 330. More particularly, in the illustrated embodiment, theattachment element 328 is located between a flange 334 of the flowdirector 318 and a flange 336 of the flow modulation insert 320. Toplace the flow modulation insert 320 in its first position 330, wherethe flow modulation insert 320 substantially blocks or restricts coolingfluid flow through apertures 324 of flow director 318, at least aportion of the attachment element 328 is raised along the radialdirection R from flange 334 of flow director 318. As such, the flowmodulation insert 320 is raised along the radial direction R, whichshifts the apertures of 326 of insert 320 out of alignment withapertures 324 of flow director 318 and thereby restricts the flow ofcooling fluid F through apertures 324. To place the flow modulationinsert 320 in its second position 332, where the flow modulation insert320 permits a greater fluid flow through apertures 324 than in the firstposition, the raised portion of attachment element 328 lowers along theradial direction R, e.g., to rest against flange 334 of flow director318, such that the apertures 326 of insert 320 substantially are alignedwith the apertures 324 of flow director 318. At the least, the apertures326 of flow modulation insert 320 in its second position 332 are more inline with the apertures 324 of flow director 318 such that the flowmodulation insert 320 does not restrict the aperture flow area to theextent the flow area is restricted when insert 320 is in its firstposition 330.

Turning now to FIG. 7 , an axial cross-section view is provided ofairfoil 300, according to another exemplary embodiment of the presentsubject matter. In the depicted embodiment, a cavity insert 338 ispositioned within cavity 316, and the flow modulation insert 320 ispositioned adjacent the inner surface 314 of body 302 of airfoil 300.Thus, space 322 is defined between the flow modulation insert 320 andcavity insert 338.

Referring to FIGS. 8A and 8B, radial cross-section views are provided ofthe airfoil 300 of FIG. 7 , according to an exemplary embodiment of thepresent subject matter. As shown in FIGS. 8A and 8B, the body 302 ofairfoil 300 defines a plurality of apertures 340 therethrough, e.g., fordirecting the flow of cooling fluid received in cavity 316 to the outersurface 312 of body 302 to provide film cooling of the outer surface.Each aperture 340 has an aperture flow area over which cooling fluidflow may flow toward outer surface 312 of body 302. The flow modulationinsert 320 defines the plurality of apertures 326 therethrough, whichmay allow the flow of cooling fluid F to flow through flow modulationinsert 320.

As described with respect to the embodiment of FIGS. 6A and 6B, in theembodiment shown in FIGS. 8A and 8B, flow modulation insert 320 changespositions in response to temperature changes to modulate the flow ofcooling fluid to body 302. More specifically, FIG. 8A illustrates thefirst position 330 of flow modulation insert 320, which the insertmaintains until a threshold temperature is reached. In the firstposition 330, flow modulation insert 320 restricts or blocks the flow ofcooling fluid to apertures 340 of body 302, i.e., flow modulation insert320 in the first position reduces the aperture flow area of eachaperture 340. In some embodiments, the first position 330 of flowmodulation insert 320 may provide a reduction in aperture flow area ofabout 50% to about 90%. In an exemplary embodiment, the first position330 reduces the aperture flow area by about 80% to about 85%. In otherembodiments, the first position 330 may block essentially all coolingfluid flow through apertures 340, i.e., the first position may providean approximately 100% reduction in aperture flow area.

At the threshold temperature, the flow modulation insert 320 transitionsto a second position 332, shown in FIG. 8B. In the second position 332,flow modulation insert 320 provides little to no restriction of coolingfluid flow through apertures 340 of airfoil body 302. That is, if theflow modulation insert 320 transitions from the first position 330 tothe second position 332, the flow modulation insert 320 opens up theapertures 340 to permit a greater flow of cooling fluid F throughapertures 340 than was permitted when the flow modulation insert 320 wasin the first position 330. In some embodiments, the second position 332of flow modulation insert 320 may block only about 5% to about 20% ofthe aperture flow area. In an exemplary embodiment, the second position332 reduces the aperture flow area by about 10% or less. In otherembodiments, the second position 332 may essentially provide norestriction of cooling fluid flow through apertures 340 of body 302,i.e., the second position may permit flow over approximately 100% of theflow area of apertures 340. Accordingly, the flow modulation insert 320may shift positions such that the location of its apertures 326 varieswith respect to the apertures 340 of body 302 to modulate the flow ofcooling fluid to body 302.

As previously described, in some embodiments flow modulation insert 320may be made from a material with a large coefficient of thermalexpansion, i.e., a high alpha material, or a shape memory alloy suchthat the flow modulation insert 320 can expand and contract in responseto temperature changes. In an example embodiment in which the flowmodulation insert 320 is formed from a high alpha material, the insert320 may be formed from a metal material and the airfoil body 302 from aCMC material such that the insert 320 expands or grows to a greaterdegree than the body 302 as the temperature rises. For example, as thetemperature increases, the apertures 326 of flow modulation insert 320may expand in cross-sectional area more than apertures 340 of body 302such that the flow modulation insert 320 no longer blocks apertures 340or restricts flow through apertures 340 to a lesser extent than theinsert 320 restricts the flow at lower temperatures. In otherembodiments, flow modulation insert 320 may be formed from a shapememory alloy such that the insert 320 shifts position with respect tobody 302 at an increased temperature to permit greater fluid flowthrough apertures 340 of body 302 and thereby provide greater cooling tothe outer surface 312 of the body 302.

Alternatively, similar to the embodiment of FIGS. 6A and 6B, rather thanforming flow modulation insert 320 from a high alpha material or a shapememory alloy, the airfoil 300 may include an attachment element 328coupled to the flow modulation insert 320 that changes position inresponse to changes in temperature and thereby shifts the position ofinsert 320 in response to changes in temperature. As shown in FIG. 8A,the attachment element 328 takes on a first configuration to place theflow modulation insert 320 in the first position 330. More particularly,in the illustrated embodiment, the attachment element 328 is locatedbetween a flange 342 of the cavity insert 338 and the flange 336 of theflow modulation insert 320. To place the flow modulation insert 320 inits first position 330, where the flow modulation insert 320substantially blocks or restricts cooling fluid flow through apertures340 of body 302, at least a portion of the attachment element 328 israised along the radial direction R, e.g., to rest against flange 342 ofcavity insert 338. As such, the flow modulation insert 320 is raisedalong the radial direction R, which shifts the apertures of 326 ofinsert 320 out of alignment with apertures 340 of body 302 and therebyrestricts the flow of cooling fluid F through apertures 340. To placethe flow modulation insert 320 in its second position 332, where theflow modulation insert 320 permits a greater fluid flow throughapertures 340 than in the first position, the raised portion ofattachment element 328 lowers along the radial direction R such that theapertures 326 of insert 320 substantially are aligned with the apertures340 of body 302. At the least, the apertures 326 of flow modulationinsert 320 in its second position 332 are more in line with theapertures 340 of body 302 such that the flow modulation insert 320 doesnot restrict the aperture flow area of apertures 340 to the extent theflow area is restricted when insert 320 is in its first position 330.

As previously stated, it will be appreciated that in embodiments such asthose illustrated in FIGS. 6A and 6B and FIGS. 8A and 8B that utilizethe attachment element 328, the attachment element 328 rather than theflow modulation insert 320 may be formed from a high alpha material or ashape memory alloy. As such, the use of the high alpha material or shapememory alloy may be limited to the attachment element 328, which may bemore economical or otherwise more efficient or beneficial than formingthe flow modulation insert 320 from such materials.

Accordingly, as described herein, a gas turbine engine may be providedwith one or more features for modulating a flow of cooling fluidtherethrough, e.g., to provide different cooling flows for differentoperating conditions or to provide a flow of cooling fluid during somebut not all operating conditions. As an example, the first position ofinserts 208, 320 may be a “cold” position, where the insert 208 or 320blocks a flow of cooling fluid when an operating temperature of a gasturbine engine is below a certain temperature. Further, the secondposition of inserts 208, 320 may be a “hot” position, wherein the insert208 or 320 unblocks the flow of cooling fluid when the operatingtemperature is at or above the certain temperature. In some embodiments,inserts 208, 320 may move between the first position and the secondposition based on changes in pressure rather than changes intemperature, e.g., the inserts 208, 320 may be in the first position atlower pressures and in the second position at higher pressures. Asanother example, insert 208 and/or flow modulation insert 320 maycomprise a shape memory alloy and have a martensite configuration and anaustenite configuration. The martensite configuration of insert 208 orinsert 320 may correspond to a first position, and the austeniteconfiguration may correspond to a second position, where the firstposition blocks a flow of cooling fluid and the second position unblocksthe flow of cooling fluid. In alternative embodiments, inserts 208, 320may be made from a temperature based shape memory alloy or from apressure based shape memory alloy. Of course, other configurations ofinsert 208 and/or flow modulation insert 320 may be used as well.

Further, in various embodiments, a gas turbine engine includes a coolingcircuit where a flow of cooling fluid directed through the coolingcircuit is modulated by throttling the power provided to the engine. Forexample, when the engine is throttled back, i.e., power to the engine isdecreased, the temperature of the cooling fluid decreases and, thus,causes a flow modulation insert with a relatively higher coefficient ofthermal expansion positioned in the cooling circuit to be exposed to acooler or lower temperature of cooling fluid. As such, the modulationinsert deflects toward a closed position to restrict the flow of coolingfluid through the cooling circuit. Conversely, when the engine isthrottled up, i.e., power to the engine is increased, the temperature ofthe cooling fluid increases, thereby causing the modulation insert to beexposed to a hotter or higher temperature of cooling fluid. Accordingly,the modulation insert deflects toward an open position to allow agreater or more cooling fluid flow through the cooling circuit. It willbe appreciated that the closed position may correspond to the firstposition and the open position may correspond to the second position, asdescribed above.

In some embodiments, modulating the flow of cooling fluid based onthrottling changes of the engine may be temperature based or, in otherembodiments, may be pressure based. That is, throttling changes of theengine may modulate the cooling flow based on a thermal change in thecooling fluid or based on a pressure change of the cooling fluid. Assuch, throttling changes of the engine may be linked to cooling circuitmodulation by a thermal mechanism or a pressure mechanism. Thus, asanother example, as the engine is throttled back the pressure of thecooling fluid is reduced and the change in pressure causes themodulation insert to deflect to the closed position, or as the engine isthrottled up, the pressure of the cooling fluid increases and themodulation insert moves to the open position. To modulate the insertposition based on pressure changes, the modulation insert either may bemechanically linked, e.g., to a pressure sensitive fulcrum or the like,or may be made from or coupled to an attachment element made from apressure sensitive shape memory alloy.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. A cooling circuit for a gas turbine engine,comprising: a cooling channel for a flow of cooling fluid, the coolingchannel defined between a radially outer surface and a radially innersurface, the radially outer surface opposite the radially inner surface,the radially inner surface exposed to a hot gas path of the gas turbineengine, the cooling channel having an inlet and an outlet definedadjacent a cooling passage; and an insert that varies in positionbetween a first position and a second position based on changes in anenvironmental temperature of the insert, wherein the first positionblocks the cooling channel to reduce the flow of cooling fluidtherethrough, wherein the second position unblocks the cooling channelto increase the flow of cooling fluid therethrough, wherein the insertis positioned over the inlet in both the first position and the secondposition such that at least a first portion of the insert and a secondportion of the insert opposite the first portion remain in a samelocation in both the first position and the second position, and whereinthe first portion of the insert is disposed on one side of the inlet ofthe cooling channel and the second portion of the insert is disposed onan opposite side of the inlet of the cooling channel.
 2. The coolingcircuit of claim 1, wherein the inlet has an inlet flow area, andwherein the first position of the insert reduces the inlet flow area byat least 50%.
 3. The cooling circuit of claim 2, wherein the firstposition of the insert reduces the inlet flow area by at least 80%. 4.The cooling circuit of claim 1, wherein the cooling channel is definedin a shroud, wherein the radially inner surface is positioned radiallyadjacent a plurality of turbine rotor blades, and wherein the inlet ofthe cooling channel is defined in a forward end of the shroud.
 5. Thecooling circuit of claim 4, wherein the insert is made from a materialthat has a larger coefficient of thermal expansion than a material fromwhich the shroud is made such that the insert expands to a greaterextent than the shroud at a given temperature.
 6. The cooling circuit ofclaim 1, wherein the insert is made from a shape memory alloy.
 7. Thecooling circuit of claim 6, wherein the shape memory alloy is a two-wayshape memory alloy.
 8. A shroud assembly for a gas turbine engine,comprising: a shroud having a radially outer surface opposite a radiallyinner surface, the radially inner surface exposed to a hot gas path ofthe gas turbine engine and positioned radially adjacent a plurality ofturbine rotor blades; a cooling channel for a flow of cooling fluiddefined in the shroud, the cooling channel having an inlet defined in aforward end of the shroud and an outlet defined within the shroudadjacent an internal cooling passage of the shroud, the forward enddefined upstream of the internal cooling passage relative to the hot gaspath of the gas turbine engine; and an insert that varies in positionbetween a first position and a second position based on changes in anenvironmental temperature or pressure of the shroud assembly, whereinthe first position blocks the cooling channel to reduce the flow ofcooling fluid therethrough, wherein the second position unblocks thecooling channel to increase the flow of cooling fluid therethrough,wherein the insert surrounds the inlet in both the first position andthe second position, wherein the insert includes a first portion and asecond portion of the insert opposite the first portion, wherein atleast the first portion and the second portion of the insert remain in asame location in both the first position and the second position, andwherein the first portion of the insert is disposed on one side of theinlet of the cooling channel and the second portion of the insert isdisposed on an opposite side of the inlet of the cooling channel.